1. Field of the Invention
This invention relates to a device capable of heating fluids to temperatures over 1800.degree. F. with a specific application as a thruster for orbital positioning or artificial earth satellites and orientation, orbital positioning, or primary propulsion of space vehicles. The device disclosed is of the type that uses an enclosed electric arc to heat propellants such as, for example, hydrazine, ammonia, nitrogen, hydrogen or bipropellant. fuel to a desired temperature prior to expansion out of a rocket propulsion nozzle. This heating to attain performance enhancement provides a high propulsive specific impulse that is in excess of that obtainable by controlled chemical reaction alone or from an electrical augmentation device or engine using electrical resistance heaters.
At present, catalytic and electrically augmented (resistance-heated) thrusters are normally used during the lifetime of three-axis stabilized or spin-stabilized satellites in order to place in, to change or to maintain orbit station. For synchronous orbit satellites, this lifetime is on the order of 8-10 years. Satellite on-board propulsion is frequently required to make major and minor corrections to achieve final orbit circularization and/or orbit station. When this is accomplished through utilization of a typical chemical reaction engine, large quantities of propellant may be expended. Use of a performance-augmented engine (using electrical energy to extend the nominal chemical reaction performance level) for this function would conserve and retain more fuel for on-orbit functions. Typically, excess electric power is available on a spacecraft even during orbit/station insertion maneuvers. Correction firings are time-spaced, with off periods between firings permitting battery recharge for subsequent firings. By this augmentation process, fuel usage can be reduced by as much as 50 percent or more.
Thrusters may also be used for correcting a satellite orbit which has decayed, or for repositioning the satellite to another location or station. Such thrusters can also be used for propelling satellites which follow other satellites, or for evading tracking satellites. Another application of the performance-enhanced engine is that of changing the orbital path of a satellite in order to make ground tracking difficult or impossible. An application of this type would be for satellite maneuvering procedures undertaken for the purposes of decoying or saturating would-be tracking capabilities of adversaries.
In usage, this engine could be ground-controlled by the spacecraft operating agency, or in some instances of covert operation may, if desired, be preprogrammed for on-orbit automatic control.
The utilization of electrical augmentation of propellant propulsion performance by means of a resistance heater is well known. Prior to the advent of practical application, the use of an electric arc to heat gases was documented, and several devices were built and laboratory-tested. With the actual flight usage of electrical resistance-heat augmented propulsion, it is now feasible to further increase propulsion performance by making specific configuration adaptations of the resistance heater devices to permit changeover to an arc heater in the thruster. Following implementation of an arc heater as a practical flight device, it is possible to further expand the engine to include self-generated or externally-applied electric or magnetic fields to provide additional propellant acceleration. The prior art and configuration approaches of hypothetical arc and arc plus accelerator devices did not address critical problems of isolation and insulation of the arc from the mounting structure or the power or thermal heat transfer efficiencies required for a useful device. Nor did they address the types of propellant and flow characteristics typical with actual flight systems with reference to how they would interact with the arc heater, nor the configuration dependence that results from flight control constraints.
The present invention addresses these constraints and provides particular design features and configurations that can make use of an arc-heated augmentation engine both possible and useful.
2. Brief Description of the Prior Art
Prior art thrusters known to applicant are of three types; those using a chemical reaction energy, those using a resistance-type heater to augment propulsive performance, and laboratory models of electric arc heaters with limited adaptability for actual spacecraft use.
Liquid-propellant fueled spacecraft engines operate at performance levels limited by the chemical reaction energy of the propellant. Performance is generally maximum for steadystate operating of more than a minute and reduced for pulsing operation. For a monopropellant-fueled engine (i.e., hydrazine) either a catalyst bed or a thermal decomposer is used to initiate the exothermic reaction process. Of these processes the catalyst bed is the more common usage. The thermal decomposer is typically brought to operating temperature by means of an electrical resistance heater. These decomposers serve only to initiate the chemical reaction, but do not add to or augment the chemical performance level. To extend the performance level, high temperature electrical resistance heaters are being used to boost the chemical performance level, thereby increasing the propellant temperature prior to its expansion through a nozzle. The devices that use a resistance-type heater are limited to the operating temperatures of the heat exchanger (under approximately 3500.degree. F.).
With prior art arc-heated laboratory thrusters, the expansion nozzle also serves as the device anode. While this would be acceptable in the laboratory, it may not be useful for many flight applications. Further, the prior art commonly uses electrical energy exclusive of propellant chemical energy, thereby eliminating from use the majority of the propellants which have been flight-qualified and are at present being used. The prior art also typically postulates the use of low molecular weight gasses or liquids as the propellant to enable the arc design to operate more easily and to obtain high specific impulses. Use of this type of propellant makes propellant storage and handling more difficult and may result in adverse spacecraft interactions. Further, no attention nor provision was given to the use of a chemically reactive propellant typical of all current flight systems.
The prior art of laboratory arc thrusters recognizes the requirements to cool the electrodes and to transfer heat from the electrodes and arc to the gas. No design attention has been given until now to thermal and electrical isolation from the propellant or mounting systems or to meet a practical usage requirement to match the device voltage to that of the space vehicle.
Prior studies of propellant accelerators combined with basic arc generators also lack understanding of the requirements and design basics that are needed for flight implementation. The following prior art is known to applicant: U.S. Pat. Nos. 2,919,370, 3,056,257, 3,149,459, 3,279,177, 3,304,719, 3,359,734, 3,447,316, 3,460,758, 3,521,453, 3,618,324, 3,651,644, 3,695,041, 3,772,885, 3,871,828, 3,956,885, 4,059,415, 4,288,982, 4,305,247: British Pat. No. 749,921; Publications: "Electric Propulsion Development" by E. Stuhlinger, March 14-16, 1962; "Survey of ElectroMagnetic Accelerators for Space Propulsion" by S. Domitz, et al., March 1966; "Flow Field Characteristics and Performance Limitations of Quasi Steady Magneto Plasma Dynamic Accelerators" by M. J. Boyle, K. E. Clarke and R. G. Jahn, March, 1975.
Further, applicant is either the sole inventor or a co-inventor in the following U.S. Pat. Nos. 3,243,954, 3,309,873, 3,324,316, 3,388,291, 3,413,509, 3,449,628, 3,452,249, 3,453,469, 3,453,474, 3,453,488, 3,453,489, 3,462,622 and 3,467,885. Below, applicant discusses those patents of the above-listed patents which applicant deems to be pertinent to the subject matter of the instant patent application. Those patents not discussed herein are believed to be of only general interest.
The type and general configuration of accelerator disclosed herein is related to some of the above-listed prior art patents. In theory, it works using some of the mechanisms disclosed in U.S. Pat. Nos. 3,388,291, 3,243,954 disclosed herein was generally disclosed in U.S. Pat. No. 3,413,509. Feeding some propellant through the cathodebuffer (about 10%) and the rest of the propellant (about 90%) through the anode was also generally disclosed in U.S. Pat. No. 3,413,509. The concept of electric field reversal at a sonic point is disclosed in 3,467,885.
The invention herein includes the following features:
1. The use of the buffered cathode and the feeding of approximately 90% of the propellant thru the anode in a space propulsion engine. In U.S. Pat. No. 3,413,509, by way of contrast, these features were incorporated into a plasma confinement device.
2. Matching the operation voltage of the propulsion engine to the bus voltage available on the spacecraft using a theoretically derived equation to define the voltage in terms of engine design parameters, a concept not shown in the prior art as disclosed herein.
3. Use of a nested set of disc-cylinder shields mounted on insulators to act both as thermal radiation shields and vacuum electrical insulators, a concept also not disclosed in the prior patents.
4. The use of a theoretical relation among the discharge current, the strength of the applied magnetic field, the ionization potential levels of the propellant as well as its molecular weight, and the geometry of the engine to establish a relation between the thrust and the required mass flow rate of propellant for stable and efficient operation of the accelerator, which relation is not previously disclosed in the above patents.
5. An anode configuration previously undisclosed in the prior patents and specifically designed to give a high mass utilization of the propellants and hence make a more efficient space propulsion engine.